The present invention generally relates to processes for producing airfoil components of turbomachinery and airfoil components produced thereby. More particularly, this invention is directed to processes for producing ceramic-based airfoil components with tip caps, and airfoil components produced thereby.
Components of turbomachinery, including blades (buckets) and vanes (nozzles) of gas turbines, are typically formed of nickel-, cobalt- or iron-base superalloys with desirable mechanical and environmental properties for turbine operating temperatures and conditions. Because the efficiency of a gas turbine is dependent on its operating temperatures, there is a demand for components that are capable of withstanding increasingly higher temperatures. As the maximum local temperature of a component approaches the melting temperature of its alloy, forced air cooling becomes necessary. For this reason, airfoils of gas turbines, and in particular their low pressure and high pressure turbine (LPT and HPT) blades, often require complex cooling schemes in which air is forced through internal cooling passages within the airfoil and then discharged through cooling holes at the airfoil surface. Airfoil components can be equipped with tip caps that regulate internal cavity pressure, allowing for proper air flow through the cooling passages and holes. Tip caps are typically cast, brazed or welded onto metallic air-cooled LPT and HPT blades.
As higher operating temperatures for gas turbines are continuously sought in order to increase their efficiency, alternative materials have been investigated. Ceramic-based materials are a notable example because their high temperature capabilities significantly reduce cooling air requirements. As used herein, ceramic-based materials encompass homogeneous (monolithic) ceramic materials as well as ceramic matrix composite (CMC) materials. CMC materials generally comprise a ceramic fiber reinforcement material embedded in a ceramic matrix material. The reinforcement material may be discontinuous short fibers that are randomly dispersed in the matrix material or continuous fibers or fiber bundles oriented within the matrix material. The reinforcement material serves as the load-bearing constituent of the CMC in the event of a matrix crack. In turn, the ceramic matrix protects the reinforcement material, maintains the orientation of its fibers, and serves to dissipate loads to the reinforcement material. Silicon-based composites, such as silicon carbide (SiC) as the matrix and/or reinforcement material, have become of particular interest to high-temperature components of gas turbines, including aircraft gas turbine engines and land-based gas turbine engines used in the power-generating industry. SiC fibers have also been used as a reinforcement material for a variety of other ceramic matrix materials, including TiC, Si3N4, and Al2O3. Continuous fiber reinforced ceramic composites (CFCC) are a particular type of CMC that offers light weight, high strength, and high stiffness for a variety of high temperature load-bearing applications, including shrouds, combustor liners, vanes (nozzles), blades (buckets), and other high-temperature components of gas turbines. A notable example of a CFCC material developed by the General Electric Company under the name HiPerComp® contains continuous silicon carbide fibers in a matrix of silicon carbide and elemental silicon or a silicon alloy.
Various techniques may be employed in the fabrication of CMC components, including chemical vapor infiltration (CVI) and melt infiltration (MI). These fabrication techniques have been used in combination with tooling or dies to produce near-net-shape articles through processes that include the application of heat and chemical processes at various processing stages. Examples of such processes, particularly for SiC/Si—SiC (fiber/matrix) CFCC materials, are disclosed in U.S. Pat. Nos. 5,015,540, 5,330,854, 5,336,350, 5,628,938, 6,024,898, 6,258,737, 6,403,158, and 6,503,441, and U.S. Patent Application Publication No. 2004/0067316. One such process entails the fabrication of CMCs from prepregs, each in the form of a tape-like structure comprising the desired reinforcement material, a precursor of the CMC matrix material, and one or more binders. After partially drying and, if appropriate, partially curing the binders (B-staging), the resulting tape is laid-up with other tapes, debulked and, if appropriate, cured while subjected to elevated pressures and temperatures to produce a cured preform. The preform is then fired (pyrolized) in a vacuum or inert atmosphere to remove solvents, decompose the binders, and convert the precursor to the desired ceramic matrix material, yielding a porous preform that is ready for melt infiltration. During melt infiltration, molten silicon and/or a silicon alloy is typically infiltrated into the porosity of the preform, where it fills the porosity and may react with carbon to form additional silicon carbide.
For purposes of discussion, a low pressure turbine (LPT) blade 10 of a gas turbine engine is represented in FIG. 1. The blade 10 is an example of a component that can be produced from ceramic-based materials, including CMC materials. The blade 10 is generally represented as being of a known type and adapted for mounting to a disk or rotor (not shown) within the turbine section of an aircraft gas turbine engine. For this reason, the blade 10 is represented as including a dovetail 12 for anchoring the blade 10 to a turbine disk by interlocking with a complementary dovetail slot formed in the circumference of the disk. As represented in FIG. 1, the interlocking features comprise one or more protrusions 14 that engage recesses defined by the dovetail slot. The blade 10 is further shown as having a platform 16 that separates an airfoil 18 from a shank 20 on which the dovetail 12 is defined.
Current state-of-the-art approaches for fabricating ceramic-based turbine blades have involved integrating the dovetail 12, platform 16, and airfoil 18 as one piece during the manufacturing process, much like conventional investment casting techniques currently used to make metallic blades. Because of their relatively higher temperature capability, CMC airfoils such as the blade 10 have not been equipped with tip caps for the purpose described above for metallic airfoil components. Moreover, brazing and welding techniques used to attach tip caps to metallic air-cooled LPT and HPT blades processes are not generally practical for attaching tip caps to airfoil components formed of CMC materials. In addition, tip caps define a geometric feature that is oriented transverse to the span-wise direction of the blade 10, such that the incorporation of a tip cap into a CMC blade would pose design and manufacturing challenges. Furthermore, the low strain-to-failure capabilities of typical CMC materials pose additional challenges to implementing tip caps in rotating CMC airfoil components such as turbine blades, where a tip cap would be subjected to high centrifugal forces.